Rotary wing vehicle

ABSTRACT

A rotary wing vehicle includes a body structure having an elongated tubular backbone or core, and a counter-rotating coaxial rotor system having rotors with each rotor having a separate motor to drive the rotors about a common rotor axis of rotation. The rotor system is used to move the rotary wing vehicle in directional flight.

This application claims priority to U.S. Provisional Application No.60/562,081 which was filed Apr. 14, 2004 and is hereby incorporated byreference herein.

BACKGROUND

The present disclosure relates to aerial vehicles and particularly tounmanned aerial vehicles (UAV). More particularly, the presentdisclosure relates to unmanned rotary wing vehicles.

Rotary wing vehicles are used in a variety of applications. Unmannedrotary wing vehicles are often used by the military, law enforcementagencies, and commercial activities for aerial reconnaissanceoperations.

SUMMARY

A rotary wing vehicle, in accordance with the present disclosureincludes a body structure having an elongated tubular backbone or core,and a counter-rotating coaxial rotor system having rotors with eachrotor having a separate motor to drive the rotors about a common rotoraxis of rotation. A power source comprising, for example, a battery,fuel cell, or hybrid gas-electric generator is provided to supplyelectric power to the motors. Power transmission to and between therotor systems is accomplished primarily by means of electrical wiringinstead of mechanical shafting. A modular structure is described whichassists manufacturability.

One embodiment of the disclosure includes an auxiliary power-pack whichis separable from the vehicle in flight to facilitate, for instance,delivery of the vehicle to a distant location. In another embodiment,the power-pack comprises a payload such as an explosive munition,dipping sonar, hydrophones, or a separable sonobouy module. Whileaspects of the disclosure are applicable to many helicopters, includingfull-sized man carrying helicopters, the current disclosure isespecially well suited for application to small, autonomous, orradio-controlled rotary wing aircraft known as remotely piloted vehicles(RPVs), or unmanned aerial vehicles (UAVs).

Additional features of the present disclosure will become apparent tothose skilled in the art upon consideration of the following detaileddescription of illustrative embodiments exemplifying the best mode ofcarrying out the disclosure as presently perceived.

BRIEF DESCRIPTION OF THE DRAWINGS

The detailed description particularly refers to the accompanying figuresin which:

FIG. 1 is a diagrammatic view of a rotary wing vehicle in accordancewith the present disclosure showing an aircraft including a guidancesystem, and a pair of rotor systems coupled to an airframe comprising anon-rotating structural spine or backbone and carrying a payload;

FIG. 2A is a perspective view of a rotary wing vehicle in accordancewith the present disclosure showing a counter-rotating coaxial rotorsystem in a vertical flight mode;

FIG. 2B is a perspective view of the rotary wing vehicle of FIG. 2Ahaving a counter-rotating coaxial rotor system and a fixed-wing boostermodule in a horizontal flight mode;

FIG. 3 is a side elevation view of the rotary wing vehicle of FIG. 2Ashowing exterior body panels, electrical wiring, and booster sectionremoved for clarity;

FIG. 4 is a side elevation view, with portions broken away, of thevehicle of FIG. 2A showing a counter-rotating coaxial rotor system andan electrical power source;

FIG. 5 is an enlarged perspective view of the vehicle of FIG. 2A, withportions broken away, showing an upper interior section of the vehicleand the counter-rotating coaxial rotor system;

FIG. 6 is an enlarged perspective view of the vehicle of FIG. 2A, withportions broken away, showing a lower interior section of the vehicleand the counter-rotating coaxial rotor system;

FIG. 7A is a perspective view of a core tube or backbone having acircular cross section and a hollow interior channel that is used as aconduit between sections of the vehicle and showing electrical wiringrunning through the hollow interior and entering and exiting at variouspoints;

FIG. 7B is a perspective view of backbone having a generally cruciformcross section with exterior channels running the length of the backbonethat can be used as conduits between sections of the vehicle.

FIG. 8 is an enlarged perspective view of a first ring mount;

FIG. 9 is an exploded perspective view of a second ring mount showingattached linkages and body supports;

FIG. 10 is an enlarged perspective view of a middle interior section ofthe vehicle of FIG. 2A, with portions broken away, showing thecounter-rotating coaxial rotor system;

FIG. 11A is an exploded perspective view of a rotor module having rotorblades with variable cyclic pitch and fixed collective pitch;

FIG. 11B is an exploded perspective view of a rotor module having rotorblades with variable cyclic and variable collective pitch;

FIGS. 12A and 12B are perspective views of a first side and a secondside of a motor mount;

FIGS. 13A and 13B are perspective views of a first side and a secondside of a rotor hub;

FIG. 14 is a sectional view taken along lines 14-14 of FIG. 2B, showingthe rotor module;

FIG. 15 is a side elevation view of the counter-rotating coaxial rotorsystem of FIG. 2A, and a core tube depending from the rotor system;

FIGS. 16A and 16B are exploded perspective views of a single powermodule including several batteries;

FIG. 17 is an orthographic view of the booster module of FIG. 2B showingone wing folded for storage and one wing extended in a flightconfiguration;

FIG. 18 is an orthographic view depicting the booster module separatingin flight from the rotary wing vehicle;

FIG. 19 is an elevation view of the rotary wing vehicle showing adipping sonar or hydrophone assembly depending from a bottom portion ofthe vehicle;

FIGS. 20A, 20B, and 20C are sequential views of the rotary wing vehicleshowing the operation of unequal length folding blades during a crashlanding of the vehicle on ground underlying the rotary wing vehicle;

FIGS. 21A and 21B are side elevation views of a storage tube and therotary wing vehicle showing the vehicle folded for storage;

FIG. 22 is a perspective view of a rotary wing vehicle in accordancewith present disclosure delivering a sensor or marking to a remotelocation shown for the purpose of illustration to be a ship on the openocean;

FIG. 23 is a side elevation view of a rotary wing vehicle folded forstorage in a rear portion of a gravity-delivered bomb; and

FIG. 24 is a perspective view of a rotary wing vehicle deploying fromthe rear of a gravity-delivered bomb to the vicinity of a target siteshowing the gravity-delivered bomb ejecting the rotary wing vehicle andthe rotary wing vehicle deploying into a vertical flight mode to loiterin the target area to provide an attacking force with real-time battledamage assessment after the gravity delivered bomb has struck thetarget.

FIG. 25A is a diagrammatic view of another rotary wing vehicle showingan aircraft having a central buss architecture with power and signalconduits, a guidance system, and a pair of rotor systems coupled to anairframe comprising a non-rotating structural spine or backbone andcarrying a payload; and

FIG. 25 B is a diagrammatic view of the rotary wing vehicle of FIG. 25Ashowing a rotor system, control system, and power supply communicatingthrough a central data/power buss with power and signal conduit.

DETAILED DESCRIPTION

As suggested diagramatically in FIG. 1, a rotary wing vehicle 1includes, in series, a first module 2, a first and a second rotor system3 and 5, power modules 13 and 14, and a second module 15 coupled inspaced-apart relation to an airframe 40 extending along a common axis 7.Illustratively, airframe 40 is an elongated central backbone and can bearranged as a hollow core or having a cruciform cross-section. Inoperation, first rotor 3 and second rotor 5 rotate in oppositedirections about common axis 7 to direct thrust in direction 24 andcreate lift in direction 24′ to cause controlled flight of rotary wingvehicle 1, as suggested in FIG. 2A. First module 2 is adapted to includea variety of guidance systems 50′, electronics 55, or payloads 15′.Second module 15 is adapted to include payload 15′, or in someembodiments, a variety of guidance systems 50′ and electronics systems55′. Payload 15′ may include, but is not limited to, munitions,radiation sensors, chemical detection sensors, biological agent sensors,active and passive listening devices, video sensors, supplemental powersources, or other mission-specific equipment. Rotary wing vehicle 1 thusprovides means for moving reconnaissance, observation, or surveymonitoring equipment to an area of interest to obtain informationtherefrom.

As suggested in FIGS. 1, 25A and 25B, first rotor system 3 includes afirst motor 54, first rotor blades 20, and a first pitch controller 56.In illustrative embodiments, motor 54 is an electric motor as shown, forexample, in FIGS. 4-6, or other suitable means for providing power torotate rotor blades 20 about common axis 7. First rotor system 3 andsecond rotor system 5 are similar to one another in structure andfunction. Second rotor system 5 includes a second motor 61, second rotorblades 22, and a second pitch controller 57. In illustrativeembodiments, motor 61 is an electric motor as shown, for example, inFIGS. 4-6, or other suitable means for providing power to rotate rotorblades 22 about common axis 7. Illustratively, electrical and electroniccomponents are connected and communicate through electrical conduit 173and electronic conduit 174 which hold power and signal lines,respectively. Although rotary wing vehicle 1 is illustrated having tworotor systems, rotary wing vehicle 1 may have more than two rotorsystems as performance and mission demands dictate.

As shown in FIGS. 1 and 3, airframe 40 is non-rotating and forms acentral elongated hollow backbone to receive first module 2, first andsecond rotor systems 3, 5, power modules 13 and 14, and second module15. Illustratively, power modules 13 and 14 are positioned to lie inside-by-side relation to one another between second rotor system 5 andsecond module 15. Because airframe 40 is hollow power modules 13, 14 canbe connected electrically through the hollow backbone to motors 54 and61.

Illustratively, pitch controller 56 is a swashplate 56′ coupled to afore/aft servo 58 and a roll servo 59 to vary the cyclic pitch of rotorblades 20 in response to input from a controller 55. In someembodiments, swashplate 56′ is further coupled to a collective servo 98to collectively change the pitch of rotor blades 20. Likewise, pitchcontroller 57 is a swashplate 57′ coupled to a fore/aft servo 58 and aroll servo 59 to vary the cyclic pitch of rotor blades 20 in response toinput from a controller 55. In some embodiments, swashplate 57′ is alsocoupled to a collective servo 98 to collectively vary the pitch of rotorblades 20. In illustrative embodiments, controller 55 is a commandsignal controller as shown, for example, in FIG. 3, or other suitablemeans for providing a desired electrical or mechanical directionalsignal to servos 58, 59, or 98, and motors 54, 61.

Illustratively, rotary wing vehicle 1 has a fixed-pitch rotor systemhaving two servos 58, 59 for aircraft pitch (helicopter-style fore/aftcyclic input) or aircraft roll (helicopter-style right/left cyclicinput) control. Servo 98, shown in phantom in FIG. 1, can be mountedsimilarly to servos 58, 59 if collective pitch control is desired. Inembodiments having a fixed-pitch rotor system, rotor systems 3,5 areconnected to swashplates 56′, 57′ by pitch links 119. Servos 58, 59 areconnected to swashplates 56′, 57′ by links 125, 126. A feature of thepresent disclosure is that rotary wing vehicle 1 can be flown with asfew as one or two cyclic servo actuators (servo 58, 59). In a“one-servo” flight mode, differential torque of motors 54, 61 controlsyaw orientation, and servo 58 controls forward and backward flight. Withonly one cyclic servo, vehicle 1 can be flown much like an airplanehaving only rudder and elevator control. In the illustrative “two-servo”flight mode, servos 58, 59 provide fore/aft aircraft pitch andright/left aircraft roll control with differential torque of motors 54,61 providing yaw control.

In operation, rotor hubs 101 rotate in opposite directions. Servos 58,59 are controlled by onboard flight control electronics to tiltsimultaneously swashplate 56′ and swashplate 57′ which then cyclicallyvary the blade pitch angle of rotating rotor blades 20 to tilt vehicle 1in one of aircraft pitch direction 170 and aircraft roll direction 171.In another embodiment having collective pitch (see FIG. 11B), collectiveservo 98 and a third pitch link (not shown) are provided to vary theaxial location of swashplates 56′, 57′ along common axis 7 and to varythe collective pitch of rotor blades 20, 22 using electronicCollective-Cyclic Pitch Mixing (CCPM). With collective-cyclic pitchmixing servos 58, 59, and 98 tilt swashplates 56′ and 57′ in unison tovary cyclic pitch and move swashplates 56′, 57′ axially in unison alongcommon axis 7 to vary collective pitch.

The illustrative embodiment employs differential motor speed for yaw(heading) control while in a vertical flight configuration. Normally,coaxial helicopters use variable blade pitch and differential bladeangle to control yaw motions in flight. In the present disclosure,differential torque generated by operating motors 54, 61 at differentspeeds relative to the fixed body of vehicle 1 generates yaw forces tostabilize and control yaw motion (i.e. rotation about common axis 7). Inthis method, the torque (and eventually the speed) of motor 54 isincreased or decreased in response to a yaw motion of rotary wingvehicle 1 about vertical common axis 7. The torque (speed) of secondmotor 61 is adjusted automatically by an onboard computer system,contained within controller 55, in opposition to the torque (speed) offirst motor 54 to maintain constant lift so that rotary wing vehicle 1neither gains nor loses altitude.

Rotor blades 20 and 22 are coupled to rotary wing aircraft 1 andsupported for rotation by rotor hubs 101. Rotor hubs 101 are furthercoupled for pivotable movement to an internal yolk 108, as shown best inFIG. 11A. Pivot axles 109 extend through rotor hub 101 and are receivedby yolk 108. Yolk 108 is adapted to couple a pair of rotor blades to hub101 for rotation about common axis 7. Yolk 108 is further coupled to afirst end of a pair of links 119. Each link 119 is further coupled on asecond end to a perimeter edge of swashplate 56′ or 57′. Thus, yolk 118is pivoted by input from swashplate 56′, 57′ in response to linearmotion input from servos 58, 59, or 98. This pivoting motion of yolk 118in turn causes each rotor blade 20, 22 to pivot in response, thusincreasing or decreasing the rotor blade pitch of rotor blades 20, 22.

As suggested in FIGS. 2A and 2B, a rotary wing vehicle 1 includes anupper section 2′, first and second rotors 3 and 5, a middle section 4, alower section 6, first and second power source modules 13, 14, and apayload 15 arranged in spaced apart relation along common axis 7.Referring now to FIGS. 2A-4, internal mechanical and electricalcomponents within upper section 2′ and middle section 4 of vehicle 1 areenclosed by a thin-walled upper body shell 10 and a middle body shell11, respectively. A lower body shell 12 covers a portion of lowersection 6, but could be extended to cover all of lower section 6. Afeature of the present disclosure is that body shells 10, 11 areblow-molded from a plastic material such as polycarbonate or ABS, and,in conjunction with backbone 40, form a structure for rotary wingaircraft that has both a central strength component and a thin exteriorcover component that together are stiff, strong and easy to manufacture.

As shown in FIG. 3, a rotary wing aircraft 1 in accordance with thepresent disclosure has a rotor system comprising a motor 54 operablyconnected to rotor blades 20 by means of a drive train such as gears106,107 (FIG. 11). A pitch control such as a swashplate 56′ (FIG. 10) isoperably connected to rotor blades 20 to vary the cyclic and/orcollective pitch of rotor blades 20 in response to output from a servoactuator such as servos 58,59 (FIG. 3) through linkages such as links125,126 (FIG. 10). Power such as electricity from batteries (not shown)or fuel from a storage tank (not shown) in a power source module 13flows through a power conduit across rotor system and provides power tooperate controller 55, motor 54, and servos 58 and 59. Control signalsfrom controller 55 flow along a signal conduit and regulate the speed ofmotor 54 and the positioning output of servos 58 and 59. The powerconduit and signal conduit are conducted between an inflow side and anoutflow side of rotor blades 20 through channels 96 formed in thestructural spine or backbone 40 (FIGS. 7A, 7B, and 15) of vehicle 1.

In hovering flight, first rotor 3 and second rotor 5 rotate in oppositedirections about common axis 7 forcing air downward in direction 24 andlifting vehicle 1 in an upwardly direction, as suggested in FIG. 2A.First rotor 3 has rotor blades 20 configured to rotate in direction 21,and second rotor 5 has rotor blades 22 configured to rotate in direction23 about common axis 7. Because first rotor blades 20 and second rotorblades 22 are equipped with a cyclic pitch control, vehicle 1 isconfigured for directional flight in direction 25 wherein common axis 7is orientated substantially vertically.

Referring now to FIG. 2B, a second embodiment contemplated by thecurrent disclosure is depicted having a booster module 8 appended tolower section 6 at a booster interface 9. Booster module 8 contains, forexample, an auxiliary power source (not shown) to augment an internalpower source contained in power modules 13 and 14 carried in vehicle 1.Illustratively, the auxiliary power source (not shown) and power modules13 and 14 are electrical batteries 13 and 14. Booster module 8 includesleft and right wings 16, 17 to provide additional lift for vehicle 1 indirectional flight in direction 18 wherein common axis 7 is orientedsubstantially horizontally.

Airframe 40 forms a structural backbone of rotary wing vehicle 1 andgenerally runs vertically through the center of rotary wing vehicle 1from upper section 2′ to lower section 6, as shown best in FIG. 4.Illustratively, airframe 40 is a non-rotating core tube with a hollowinterior channel 96 (FIG. 7A) or a cruciform beam 97 with exteriorchannels (FIG. 7B). First and second rotor modules 3 and 5, allcomponents within upper section 2′, middle section 4, and lower section6 are coupled to airframe 40. Referring now to FIG. 7A, non-rotatinghollow core tube 40 further acts as a conduit for electrical wiring 45,plumbing (not shown), and mechanical linkages (not shown) passingbetween components in upper section 2′, middle section 4, and lowersection 6 of rotary wing vehicle 1. Longitudinal slots 46 and 47 areprovided as entry and exits points for wiring 45, plumbing, andlinkages. Since non-rotating hollow core tube 40 and cruciform beam areunitary and continuous between body sections 2, 4 and 6, the rigidityand light-weight structural properties of vehicle 1 are increased.Illustratively, non-rotating hollow core tube 40 and cruciform beam 97are preferably made of wound or pultruded carbon graphite fiber,fiberglass, or aluminum alloy number 7075 (or similar) with an outsidediameter (core tube 40) or width dimension (cruciform beam) of about 0.5inches (13 mm) and a wall thickness of between about 0.03 inches (0.76mm) and about 0.05 inches (1.3 mm).

Rotary wing vehicle 1 is arranged having three body sections, as shownbest in FIG. 3. Upper section 2′ is arranged having a horizonsensor/stabilizer 50, an electronic gyro stabilizer 51, a gyro mountingtable 52 coupled to an upper end of core tube 40, a first motor speedcontroller 53, a first motor 54, a radio receiver, and controller 55.Middle section 4 includes a first swashplate 56′, a second swashplate57′, a fore-aft cyclic servo 58, and a roll cyclic servo 59. Lowersection 6 includes a second motor speed controller 60, a second motor61, a radio battery 62, first and second battery modules 13 and 14, andpayload module 15.

In the illustrated embodiment, horizon sensor/stabilizer 50 is a model“FS8 Copilot” model by FMA company, gyro stabilizer 51 is a “G500” modelsilicone ring gyro by JR company, motors 54, 61 are “B2041S” models byHacker company, and motor speed controllers 53, 60 are “Pegasus 35”models by Castle Creations company which are computer-based digitalprogrammable speed controllers. Rotary wing vehicle 1 is also configuredto receive a GPS receiver/controller and telemetry system (not shown),arranged to be coupled to upper section 2′.

Interior components of rotary wing vehicle 1 are coupled to core tube 40by ring mounts 70, as shown in FIG. 8. Ring mount 70 includes an annularinner portion 71 conforming to the annular exterior surface of core tube40. Ring mount 70 includes radially extending mounting arms 72, 73, 74having flanges 75, 76, 77 adapted to hold mechanical, electrical andother interior components of rotary wing vehicle 1. Ring mount 70 isarranged to support motor 54 in flange 75, motor speed controller 53 onflange 76, and radio receiver 55″ on flange 77. Interior components ofvehicle 1 are coupled, for example, to mounting flanges using a varietyof fasteners (such a nylon ties through apertures 78) or adhesives.Annular portion 71 provides means for locking ring mount 70 tonon-rotating hollow core tube 40 to prevent ring mount 70 from rotatingor sliding axially along non-rotating hollow core tube 40. Means forlocking ring mount 70 to non-rotating hollow core tube 40 includesfasteners (not shown) received by set screw receiver 79 or a variety ofadhesives. A second ring mount 80, as shown in FIG. 9, includes anannular ring 81, arms 82 and 83, and axial posts 84, 85 for supportingbody standoffs 86, 87, 88, swashplate anti-rotation arms 90 and 91, andswashplate links 92 and 93.

Servo module 81 includes ring mount 80 supporting pitch servo 58, rollservo 59, and universal body standoffs 86, 87 (as described in U.S.Provisional Patent Application No. 60/525,585 to Arlton which is herebyincorporated by reference herein) which support middle body shell 11, asshown, for example, in FIG. 10. Ring mounts 70, 80 are arranged toincorporate and support many structural features of rotary wing vehicle1. Ring mounts 70, 80 assist assembly of rotary wing vehicle 1 becausering mounts 70, 80 and associated interior components can bepreassembled as subassemblies and then later assembled along with othermodules to non-rotating hollow core tube 40 in a final manufacturingstep.

Referring now to FIGS. 11A, 12A, 12B, 13A, 13B and 14, rotor module 3includes a rotor mount 100, a rotor hub 101 having an internal gear 107,first and second ball bearings 102 and 103, a ring clip 104, motor 54, aplanetary gearbox 105, a pinion gear 106, a blade yolk 108, pivot axles109, axle end caps 110, torsion springs 111, and rotor blades 20. Amotor mount 122 is receptive to gearbox 105 to couple motor 54 to rotormount 100. When assembled, bearings 102, 103 are retained by ring clip104 engaging slot 108 on a boss 112 extending from rotor mount 100.Blade 20 is held in place by a pin 113 extending through cap 110 andaperture 114 formed in axle 109. Axle 109 passes through a bearingaperture 117 formed in hub 101 and into an aperture 118 in yolk 108 whenit is retained by another pin (not shown). Links 119 couple yolk 108 toswashplate 56′.

As shown in FIG. 11B, a rotor module adapted to support both cyclicallyand collectively pitchable rotor blades includes collective rotor hub201 that is similar to hub 101 and receptive to a collective yolk frame208 coupled to bosses 214 formed on an interior surface of hub 201 byfasteners 212. Collective yolk frame 208 supports the radial flightloads produced by rotor blades 20 acting through thrust bearings 203.Links 119 couple pitch arms 210 to swashplate 56′.

Illustratively, planetary gearbox 105 has a reducing speed ratio ofabout 4:1. Pinion gear on motor 54 has nine teeth and engages internalgear 107 on rotor hub 101 which has sixty teeth, so the total speedreduction ratio of rotor module 3 is about 26.7:1 (that is, the outputshaft of motor 54 turns 26.7 times for each turn of rotor hub 101). Thisreduction ratio encourages the use of high efficiency electric motorsrunning at high voltages and high speeds.

Illustratively, motor 54 is a brushless motor. In some applications,especially where flight times are short and economy is a factor (forexample, in a short-range disposable munition) several low-cost brushedmotors (i.e. motors having carbon brushes and rotating commutators) areused in place of one high-cost brushless motor 54 to turn rotor hub 101.In such cases, while rotor module 3 is shown having one motor 54 todrive rotor hub 101, it is within the scope of this disclosure toinclude several motors around the circumference of rotor mount 100 todrive rotor hub 101 instead of only one. It is also anticipated thatrotor hub 100 itself can be configured with wire coils and magnets toact as a motor so that no separate motors are required to drive rotorhub 101 about common axis 7.

Rotor blade 20 in the embodiment shown is injection molded ofpolycarbonate plastic material and is of the type described in U.S. Pat.No. 5,879,131 by Arlton, which patent is hereby incorporated byreference herein. Rotor blade 20 is free to flap upward and downwardabout 6 degrees about flapping axis 120 before tabs 121 on torsionsprings 111 contact pitch axle 109 and resist further flapping. Thismeans that rotor blades 20 can flap up and down freely in flight about+/−6 degrees and can fold upward 90 degrees and downward 90 degrees forstorage or during a crash landing.

In the embodiment shown in the drawings, rotor mount 100 is injectionmolded in one piece from a thermoplastic material such as polycarbonateor nylon. Rotor hub 101 is injection molded in one piece from athermoplastic material such as nylon or acetal. Rotor blades 20 aresupported in flight by rotor hub 101 (which forms part of the exteriorbody shell of vehicle 1) instead of by traditional coaxial shaftscoincident with common axis 7. This places rotor support bearings 103,104 very close to rotor blades 20 and frees space within the centralbody portion of rotary wing vehicle 1 for other mechanical or electricalcomponents. In a fixed-pitch rotor system (shown in the drawings) radialflight forces produced by rotating blades 20 are supported by internalyolk 108 which connects two rotor blades 20 and which includes aninternal aperture surrounding and bypassing core tube 40, thus nospecial thrust bearings are required.

Referring now to FIG. 15, a coaxial rotor system in accordance with thecurrent disclosure comprises core tube 40, two rotor systems 3, 5, twoswashplates 56′ and 57′, and one servo module 81 coupled to non-rotatinghollow core tube 40 in mirrored symmetry around servo module 81. While acoaxial rotor system with two rotors is disclosed, rotary wing vehicle 1could be equipped with additional rotor systems (not shown) spaced apartalong the length of non-rotating hollow core tube 40 for additionalthrust or operational capabilities.

In the illustrated embodiment, rotary wing vehicle 1 has a fixed-pitchrotor system which requires only two servos 58, 59 for aircraft pitch(fore-aft cyclic) and aircraft roll (right-left cyclic) control. A thirdcollective servo 98 can be mounted in a similar fashion in middlesection 4, for instance, if collective pitch control is desired.

Rotor systems 3,5 are connected to swashplates 56′, 57′ by pitch links119. Servos 58, 59 are connected to swashplates 56′, 57′ by links 125,126. In operation, rotor hubs 101 rotate in opposite directions. Servos58, 59 are controlled by onboard flight control electronics 55′ to tiltsimultaneously swashplate 56′ and swashplate 57′ which then cyclicallyvary the blade pitch angle of rotating rotor blades 20 to tilt vehicle 1in one of aircraft pitch direction and aircraft roll direction. Inanother embodiment having collective pitch (see FIG. 11B), a third servoand third pitch link (not shown) are provided to vary the axial locationof swashplates 56′, 57′ along common axis 7 and to vary the collectivepitch of rotor blades 20, 22 using electronic Collective-Cyclic PitchMixing (CCPM). Using servos positioned to lie between rotor systems 3, 5and directly coupling control swashplates 56′, 57′ with linkages tocontrol a coaxial rotor system in this way is a feature of theembodiment.

An illustrative feature of the disclosure is that motors 54, 61 arepositioned to lie on opposite sides of (above and below) rotors 3, 5with power transmission between the rotors accomplished throughelectrical wiring 45 instead of mechanical shafting thereby reducingmechanical complexity and weight. In another embodiment (not shown),motors 54, 61 are positioned to lie between the rotors, and servoactuators 58, 59 are positioned to lie in spaced-apart relation tolocate rotors 3, 5 therebetween. Because power and control of the rotorsystem is entirely electrical in nature, the entire control system ofrotary wing vehicle 1 can be operated electrically by digital computersand solid-state electronics without mechanical linkages or hydraulicamplification. Locating two sets of motors on opposite sides of rotors3, 5, and on opposite sides of servo module 81 eliminates the need forconcentric rotating shafting between rotors 3, 5, and positions servos58, 59 to drive both swashplates 56′, 57′ directly.

A feature of the present disclosure is that vehicle 1 can be flown withas few as one or two cyclic servo actuators (servo 58, 59). In aone-servo flight mode, differential torque of motors 54, 61 controls yaworientation, and servo 58 controls forward and backward flight. Withonly one cyclic servo, vehicle 1 can be flown much like an airplanehaving only rudder and elevator control. In a two-servo flight mode, asillustrated in the drawings, servos 58, 59 provide fore/aft aircraftpitch and right/left aircraft roll control with differential torque ofmotors 54, 61 providing yaw control.

In another embodiment of the current disclosure, power to drive motors54, 61 in flight is provided by high-capacity electric batteries 130such as lithium-polymer or lithium-ion batteries, or fuel cells.Referring now to FIGS. 16A and 16B, power module 13 has six rechargeablelithium ion batteries 130 arranged in a hexagonal pattern aroundnon-rotating hollow core tube 40 and wired in series to produce about21.6 volts of electrical potential. Battery ring mount 131 is formed toinclude center aperture (ring) 132 to accommodate non-rotating hollowcore tube 40 and flange 133 to hold batteries 130. Power wires 45 (notshown) from battery module 13 enter non-rotating hollow core tube 40 atopening 47 (see FIG. 7), and are routed through non-rotating hollow coretube 40 to motor speed controllers 53, 60.

As shown best in FIG. 25A multiple power modules 13, 14 are provided foradditional energy capacity during flight and are, illustratively, wiredin parallel to increase the electrical current available to motors 54,61. Flight times of rotary wing vehicle 1 can be adjusted by adjustingthe number of power modules 13, 14 carried in flight.

Extra locking rings (or ring mounts with no radial arms) 135 areprovided above and below power module 13, 14 to help couple powermodules 13, 14 to non-rotating hollow core tube 40, as shown, forexample, in FIG. 4. Since power modules 13, 14 are relatively heavycompared to other components of vehicle 1, locking rings 135 preventpower modules 13, 14 from sliding along non-rotating hollow core tube 40during a crash landing of rotary wing vehicle 1. A feature of thepresent disclosure is that rotary wing vehicle 1 is well-suited to bemanufactured and assembled in modules. Rotor, wing, control, power,booster, electronics, and payload modules are manufactured separatelyand slid onto core tube 40. Electrical connectors for connectionspassing through openings 46, 47 in core tube 40 are mounted flush withthe surface of core tube 40 to assist in assembly and disassembly ofvehicle 1 for maintenance and repairs.

Energy density and power density are considerations in UAV design andcan be applied to an aircraft as a whole. Aircraft with higher energydensities and power densities have better overall performance thanaircraft with lower densities. In general, energy density and powerdensity are defined as the amount of energy and power available per unitweight. For example, the energy density of a fuel or electric battery(also known as “specific energy”) corresponds to the amount of energycontained in a unit measure of fuel or battery (measured, for instance,in Nm/Kg or ft-lbs/slug).

Chemical (liquid) fuels tend to have higher energy densities thanelectric batteries. One additional characteristic of liquid fuel poweras compared to electric battery power is that the weight of a liquidfueled aircraft decreases over the course of a flight (as much as 60%)as it burns fuel. Consequently the energy density of a liquid fueledaircraft (i.e., the energy available per unit weight of the aircraft)decreases slowly and power density (power available per unit weight)increases as it flies. This means that the performance of liquid fueledaircraft actually improves near the end of a flight.

In contrast, the overall power density of an electric-powered aircraftis constant throughout the flight because the maximum output power ofthe batteries is almost constant and the batteries do not lose weight asthey discharge. Energy density also decreases quickly because the totalenergy available decreases. To improve energy and power density of thecurrent disclosure, an auxiliary electric booster or power module 8 isprovided that can be jettisoned in flight after its energy supply isdepleted. Thus, booster module 8 comprises additional battery modules(not shown) assembled around common axis 7 with a mechanism to retainbooster module 8 to rotary wing vehicle 1.

In another embodiment, booster 8 includes an internal combustion engine(such as a diesel engine not shown) which drives an electric generator(not shown) to convert chemical energy contained in a chemical fuel toelectrical energy. In other embodiments contemplated by this disclosure,a turbo-electric generator system (not shown) may be used to createelectrical energy. A consideration of a booster module 8 containing sucha gas-electric generator is that the entire weight of the module, fuelsystem, and engine, can be jettisoned at the end of a first flight phaseleaving the relatively low weight rotary wing vehicle 1 to complete asecond flight phase.

In the illustrative embodiment, booster module 8 includes foldable wings16, 17 to increase lift in a horizontal flight mode of rotary wingvehicle 1. As shown in FIG. 17, wing 17 is folded about folding axis 140for compact storage. Wings 16, 17 are attached at about their “quarterchord” location to pivot shafts (not shown). When deployed for flightwith pivot shafts held rigidly perpendicular to common axis 7 (see alsoFIG. 2), wing 16 is free to pivot about pitch axis 143 to find its ownbest angle of attack. Because wings 16, 17 are free to rotate abouttheir own pitch axes in flight, appendages such as wings 16, 17 aresometimes referred to as “free-wings.” It should be noted that wings 16,17, being free-wings, can operate efficiently over a wide speed rangebecause of their ability to change pitch automatically to meet theoncoming airflow. Application of such a free wing to a rotary wing UAVis a feature of the disclosure.

In high-speed horizontal flight, common axis 7 is orientatedsubstantially horizontally with rotor modules 3, 5 together acting likea single counter-rotating propeller to pull rotary wing vehicle 1 in ahorizontal direction 18. Wings 16, 17 help to lift lower section 6 andbooster module 8 so that rotor modules 3, 5 can apply more power toforward propulsion and less to vertical lifting.

It should also be noted that the current disclosure does not requireaerodynamic control surfaces (such as on wings 16, 17) because cycliccontrol of rotor module 3, 5 provides control power for maneuvering inaircraft pitch (elevation) direction 144 and aircraft yaw (heading)direction 145 when common axis 7 is substantially horizontal.Airplane-style roll control (about common axis 7) during high-speedhorizontal flight is accomplished though differential torque/speed ofrotor modules 3, 5. This method of control for horizontal flight of arotary-wing UAV is a feature of the illustrative embodiment.

Referring now to FIGS. 18A and 18B, when the energy of booster module 8has been depleted, a command from on-board controller 55 of rotary wingvehicle 1 actuates a mechanism such as a latch (not shown) thatseparates booster module 8 from rotary wing vehicle 1 and booster module8 falls away in direction 19. Rotary wing vehicle 1 then, in one flightmode, assumes a more vertical orientation and flies like a helicopter.

In another embodiment, booster module 8 includes a mission-specificpayload 147 such as an explosive munition, dipping sonar, hydrophones,radio ID marker, or a sonobouy. As illustrated in FIG. 19, uponseparation from rotary wing vehicle 1, booster module 8 falls awayleaving a sonar or hydroponic system 147 or other sensor connected torotary wing vehicle 1 by wire or fiber optic cable 146 so that rotarywing vehicle 1 can move payload 147 from place to place, deliver payload147 accurately to a desired location, and act as a telemetry linkbetween payload 147 and a remote receiver (not shown). This can be aneffective method of, for example, monitoring a target or marking a shipat sea with a remote radio ID marker or other marking instrument.

FIG. 22 illustrates a method of delivering a marker comprising, forexample, a sensor, or a marking device, such as indelible paint or aradio transmitter, to a remote location, in this case a ship on an openocean 157. Vehicle 1 is shown approaching ship S (in frame), maneuveringto touch ship S and leaving the marker on ship S (in frame) and exitingthe area (in frame). This method of marking is a feature of the presentdisclosure that allows a point of interest to be monitored after vehicle1 has left the local area. Alternatively or in conjunction, vehicle 1can retain a sensor when it leaves the local area which may, forinstance, have taken a sample of the atmosphere near ship S, and returnthe sensor and sample to a remote processing point for further analysisby a mass spectrometer, biological or radiological measuring device orother such device (not shown). While the point of interest shown in thedrawings as a ship S, it will be understood that ship S could be anyother point of interest accessible to vehicle 1 such as a truck,aircraft, building, tower, power line, or open area of land.

Another embodiment of the current disclosure shown in FIGS. 20A, 20B and20C, has unequal length folding, coaxial rotor blades 148, 149 withupper blades 148 having a greater span than lower blades 149. This is afeature arranged so that during a crash landing of vehicle when upperblades 148 contact the ground 155 before lower, shorter blades 149 sothat upper blades 148 fold away from, or faster than, lower blades 149thereby reducing the possibility that upper blades 148 and lower blades149 will contact each other while still rotating at high speed. As shownin the drawings, lower blades 149 span about 20 to 22 inches (51 cm to56 cm).

The ability to fold for compact storage and for landing is anotherfeature of the current disclosure. As shown in FIGS. 21A and 21B, rotarywing vehicle 1 is compact enough to fit inside a standard A-sizesonobouy tube used by the United States Navy. The unique core-tubestructure of the current disclosure not only allows rotary wing vehicle1 to be miniaturized to fit within a sonobouy tube, it also absorbs theforces of launch with a Charge Actuated Device (CAD) from an aircraftsuch as the Navy's P-3 maritime surveillance aircraft.

In one embodiment suggested in FIG. 21A, disposable launch canister 150is provided to protect the aerodynamic surfaces of rotary wing vehicle 1as it is launched from an aircraft traveling 150-250 knots at analtitude of 10,000 to 20,000 feet. A parachute (not shown) attached tocanister 150 slows and stabilizes the descent of canister 150 whichseparates from rotary wing vehicle 1 at a lower altitude.Illustratively, rotary wing vehicle 1 is shown to scale and has a bodylength 30 of about 24 inches (51 cm), upper diameter 31 of about 2.25inches (5.7 cm), upper rotor diameter 32 of about 28 inches (71 cm) andlower rotor diameter 33 of about 24 inches (61 cm) or less. Boostermodule 8 has a length 34 of about 12 inches (30 cm). First rotor 3 andsecond rotor 5 rotate at about 1400 RPM in hovering flight and at aboutor above 2000 RPM during vertical ascent and high-speed maneuvers.

Another embodiment contemplated by this disclosure is adapted for usewith a munition for assessing target damage done by the munition. Asshown in FIG. 23, vehicle 1 is adapted for use with the munition,illustratively shown in the drawings as a gravity-delivered bomb 160.Bomb 160 is dropped from a launch platform such as an aircraft. Inoperation, gravity-delivered bomb 160 transports vehicle 1 to thevicinity of a target site whereupon vehicle 1 is released to fall awayfrom bomb 160, illustratively slowed by use of an auxiliary drag chute162, or ejected from bomb 160 by an explosive charge-actuated device,before bomb 160 reaches its target. Vehicle 1 then orbits or hovers inthe target area near the impact site to observe bomb damage andtransmits video and other information to a remote operator (not shown).This method of munition damage assessment is a feature of the disclosurewhich provides immediate battle damage assessments without requiring alaunch platform to remain in the strike zone and reduces the need forsubsequent strikes against the same target while minimizing risk tohuman crew members.

One feature of the disclosure is the non-rotating hollow core tube 40 orcruciform beam structural backbone that can, in some embodiments, doubleas a conduit for wiring and plumbing. A method or system of assemblingmechanical and electrical components to the core or backbone isdescribed to promote ease of assembly of a variety of UAVs from a kit ofbasic modules.

Another feature is that each of the rotors 20, 22 of the coaxial systemof the current disclosure are driven by one or more separate electricmotors, and the motors are positioned to lie on opposites sides of therotors, with power transmission to and between the motors accomplishedthrough electrical wiring (passing through the hollow core) instead ofmechanical shafting, clutches, and gears. Compact rotor assembliessupport the rotors for rotation without the need for traditionalrotating coaxial shafting.

Still another feature is that a swashplate control system and one ormore electric motors are provided for each rotor and are positioned tolie on opposite sides of each rotor thereby simplifying the mechanicaland electrical connections needed to drive and control the rotors. Rotormodules are provided to quickly and easily assemble systems of rotors tothe hollow core. Multiple rotor modules and swashplates are controlledby a single group of servos housed in a module.

An additional feature is that folding rotor blades 148, 149 are ofunequal length. On the current disclosure with counter-rotating rotors3, 5, folding blades 148, 149 of unequal length reduce the chance thatthe blades will contact one another as they fold at high speed during acrash-landing.

Another feature off the disclosure is a method of improving energy andpower density on UAVs which can include a booster module 8 which isseparable from the main vehicle in flight. A booster module 8 isprovided to operate the UAV during a first flight phase. At the end ofthe first flight phase, the booster module falls away thereby reducingthe weight of the UAV for continued operation in a second flight phase.On electric powered UAVs the power module can comprise a pack ofbatteries with or without an auxiliary lifting surface which isjettisoned in flight after the battery power is depleted, or payloadsspecific to a particular mission.

1. A rotary wing aircraft comprising an airframe, a first rotor systemon the airframe, the first rotor system including first rotor bladessupported by a first rotor shaft for rotation about an axis of rotation,a first pitch controller, and a first motor, and a second rotor systemon the airframe, the second rotor system including second rotor bladessupported by a second rotor shaft for rotation about the axis ofrotation, a second pitch controller, and a second motor, and wherein thefirst rotor shaft and second rotor shaft are positioned to lie inaxially spaced apart relation to one another along the axis of rotation.2. The rotary wing aircraft of claim 1, wherein the first rotor bladesare positioned to lie in spaced-apart relation to the second rotorblades and the first and second motors are positioned to lie inspaced-apart relation to one another to locate the first and secondrotor blades therebetween.
 3. The rotary wing aircraft of claim 2,further comprising a first module and a second module, wherein the firstrotor system is positioned to lie in spaced-apart relation to the secondrotor system and the first module and the second module are positionedto lie in spaced-apart relation to one another to locate the first andsecond rotor systems therebetween.
 4. The rotary wing aircraft of claim2, wherein the first pitch controller and the second pitch controllercomprise swashplate means and servo means, and are positioned to lieadjacent to one another in spaced-apart relation between the first andsecond rotor blades.
 5. The rotary wing aircraft of claim 2, wherein thefirst rotor blades are positioned to lie in spaced-apart relation to thesecond rotor blades and the first pitch controller and second pitchcontroller are positioned to lie in spaced apart relation to one anotherto locate the first and second rotor blades therebetween.
 6. The rotarywing aircraft of claim 1, wherein the first motor is positioned to liein spaced-apart relation to the second motor and the first and secondrotor blades are positioned to lie in spaced-apart relation to oneanother to locate the first and second motors therebetween.
 7. Therotary wing aircraft of claim 6, further comprising servo actuators andlinkages, wherein the first pitch controller and second pitch controllershare a common set of servos and linkages and are positioned to lie inspaced-apart relation to one another between the first rotor blades andthe second rotor blades.
 8. The rotary wing aircraft of claim 6, furthercomprising a first module and a second module, wherein the first rotorsystem is positioned to lie in spaced-apart relation to the second rotorsystem and the first module and the second module are positioned to liein spaced-apart relation to one another to locate the first and secondrotor systems therebetween.
 9. The rotary wing aircraft of claim 6,further comprising a non-rotating structural backbone connecting thefirst rotor system and the second rotor system, and a power modulecoupled to the non-rotating structural backbone, wherein energy from thepower module on one side of a rotor plane of rotation is conducted tothe motor on the opposite side of the same rotor plane of rotationthrough the structural backbone.
 10. The rotary wing aircraft of claim6, wherein the first rotor blades are positioned to lie in spaced-apartrelation to the second rotor blades and the first pitch controller andsecond pitch controller are positioned to lie in spaced-apart relationto one another to locate the first and second rotor blades therebetween.11. The rotary wing aircraft of claim 1, wherein the first rotor systemis positioned to lie in spaced-apart relation to the second rotor systemand each motor is positioned to lie in spaced-apart relation to itsassociated rotor blades to locate their associated pitch controllerstherebetween.
 12. A rotary wing aircraft comprising first and secondcounter-rotating rotors rotatable about a common rotor axis, each rotorhaving variable pitch rotor blades, a flight control system having atleast one electronic signal processing device for processing flightcontrol commands, electronic command signal means for controlling aservo actuator connected to both rotors, and at least two swashplateslinked to the servo actuator, the servo actuator being arranged tocontrol cyclic blade pitch of both rotors in unison in response to asignal provided by the electronic command signal means to the servoactuator.
 13. The rotary wing aircraft of claim 12, wherein the meansfor controlling a servo actuator connected to both rotors includes acommand signal controller, and each swashplate includes a first and asecond pitch controller connected to the first and second rotor systemsand are positioned to lie axially between the first and second rotorsystems.
 14. The rotary wing aircraft of claim 13, further comprising asecond servo actuator, and wherein the command signal controller isconnected to the first and second servos to generate electronic signalsto the first and second servo actuators to vary roll and pitch cyclic ofboth rotor systems.
 15. The rotary wing aircraft of claim 13, furthercomprising a third servo actuator, and wherein the command signalcontroller is connected to the first, second, and third servos togenerate electronic signals to the first, second, and third servoactuators to vary cyclic and collective blade pitch of both rotorsystems.
 16. The rotary wing aircraft of claim 13, further comprising afirst and second motor coupled to the first and second rotors,respectively, and wherein the command signal controller is connected tothe first and second motors to generate electronic torque controlsignals to at least one of the first motor and second motor to vary thespeed of at least one of the first and second rotors to control at leastone of total thrust and total torque of the combined rotor system. 17.The rotary wing aircraft of claim 12, further comprising a non-rotatingstructural backbone, wherein a cross section of the non-rotatingstructural backbone includes open areas arranged to extend along alength of the aircraft such that energy from a power module on one sideof a rotor plane of rotation is conducted to a motor on an opposite sideof the same rotor plane of rotation through the structural backbone. 18.The rotary wing aircraft of claim 17, further comprising a first and asecond module, and wherein the first power module and the second rotorsystem are positioned to lie in spaced-apart relation to one another tolocate the first rotor system therebetween, and the first module and thesecond power module are positioned to lie in spaced apart relation toone another to locate the first power module and the second rotor systemtherebetween.
 19. The rotary wing aircraft of claim 12, wherein thefirst and second swashplates are operably connected by swashplatelinkages to move in unison.
 20. The rotary wing aircraft of claim 19,wherein the first and second swashplates are actuated by at most twoservo actuators which tilt the first and second swashplates in unison.21. The rotary wing aircraft of claim 19, wherein the first and secondswashplates are actuated by at most three servo actuators and the threeservo actuators control cyclic blade pitch by tilting the first andsecond swashplates in unison and collective blade pitch by translatingthe first and second swashplates in unison parallel to the rotor axis.22. The rotary wing aircraft of claim 12, further comprising electroniccontrol signals for varying the speed of rotation of the first andsecond rotors, and wherein altitude control of the rotary wing aircraftis accomplished by varying the speed of the first and second rotorssubstantially in unison.
 23. The rotary wing aircraft of claim 22,further comprising electronic control signals for varying the speed ofrotation of the first and second rotors, and wherein rotational controlof the rotary wing aircraft about the parallel rotor axes isaccomplished by varying the speed of the first and second rotors inopposition to each other.
 24. The rotary wing aircraft of claim 22,wherein the axis of rotation is substantially vertical during flight.25. The rotary wing aircraft of claim 22, wherein the axis of rotationis substantially horizontal during flight.
 26. A rotary wing aircraftcomprising a fixed-wing flight structure supportive of horizontal flightin a horizontal flight mode, and a rotary-wing flight structuresupportive of vertical flight in a vertical flight mode, and wherein theaircraft is reconfigurable in flight from one flight mode to the otherflight mode by jettisoning the structure associated with the alternateflight mode.
 27. The rotary wing aircraft of claim 26, wherein theaircraft begins flight in a fixed-wing horizontal flight mode andconverts to a rotary-wing vertical flight mode by jettisoning thefixed-wing flight structure in flight.
 28. A kit for building a rotarywing aircraft comprising a structural backbone, at least one propulsionmodule including a motor-rotor system having rotor blades, a controlmodule including an electronic control system, a power module with astore of usable energy, and a payload module for carrying payloads. 29.A method of controlling the angular orientation of a fixed-wing aircrafthaving counter-rotating rotor blades, the method comprising varyingcyclic blade pitch of the rotor blades to control aircraft pitch andyaw, and varying the rotational speed of the rotor blades to producedifferential torque between the rotors to control aircraft roll.
 30. Amethod of accurately deploying a sensor to a remote location, the methodcomprising a first step of deploying an unmanned aerial vehicle in ahigh-speed horizontal flight configuration to reach a destinationlocation, a second step of jettisoning the horizontal flight structureof the unmanned aerial vehicle and reconfiguring the unmanned aerialvehicle for substantially vertical hovering flight, and a third step offlying in a vertical orientation accurately to deploy the sensor.
 31. Amethod for increasing power and energy density in UAV's in flight, themethod comprising the steps of (1) equipping the UAV with power packssuch as electrical batteries, (2) drawing power from one or more powerpacks in flight, (3) jettisoning a power pack in flight as the storedpower in the pack is depleted, and (4) repeating steps 2-3 as necessaryuntil the end of the flight.
 32. The method of claim 31, the methodfurther comprising the step of jettisoning portions of the UAV structurein flight.
 33. A rotary wing aircraft comprising an airframe, a firstrotor system on the airframe, the first rotor system including firstrotor blades supported for rotation in a first rotor plane about an axisof rotation and a first motor, and a second rotor system on theairframe, the second rotor system including second rotor bladessupported for rotation in a second rotor plane about the axis ofrotation and a second motor, and wherein the first motor is positionedto lie on an inflow side of the second rotor system and the second motoris positioned to lie on an outflow side of the first rotor system. 34.The rotary wing aircraft of claim 33, wherein at least one rotor systemincludes variable pitch rotor blades controlled by a blade pitchcontroller.
 35. The rotary wing aircraft of claim 34, wherein the atleast one rotor system rotates in a rotor plane of rotation and theblade pitch controller and motor are positioned to lie in spaced apartrelation to one another to locate the rotor plane of rotationtherebetween.
 36. The rotary wing aircraft of claim 33, wherein thefirst rotor system includes variable pitch rotor blades controlled by afirst pitch controller, the second rotor system includes variable pitchrotor blades controlled by a second pitch controller, the first motorand first pitch controller are positioned to lie in spaced-apartrelation to one another to locate the first rotor plane of rotationtherebetween, and the second motor and second pitch controller arepositioned to lie in spaced-apart relation to one another to locate thesecond rotor plane of rotation therebetween.
 37. The rotary wingaircraft of claim 36, wherein the first pitch controller and secondpitch controller comprise swashplate means and servo means connected tothe first and second rotor systems and are positioned to lie axiallybetween the first and second rotor systems.
 38. The rotary wing aircraftof claim 37, wherein the first pitch controller and second pitchcontroller share a common set of servo actuators and linkages.
 39. Therotary wing aircraft of claim 36, wherein the first motor and secondmotor are positioned to lie axially between the first and second rotors.40. The rotary wing aircraft of claim 36, further comprising anon-rotating structural backbone connecting the first rotor system andsecond rotor system, and wherein electronic control signals areconducted from one side of a rotor plane of rotation to an opposite sideof the same rotor plane of rotation through the non-rotating structuralbackbone.
 41. The rotary wing aircraft of claim 36, further comprising anon-rotating structural backbone connecting the first rotor system andsecond rotor system, and wherein energy from a power module on one sideof a rotor plane of rotation is conducted to a motor on an opposite sideof the same rotor plane of rotation through the structural backbone. 42.A rotary wing aircraft comprising an airframe having a non-rotatingstructural backbone extending along a longitudinal aircraft axis, andfirst and second rotor systems connected to the airframe and supportedfor rotation about the longitudinal axis by the structural backbone, andwherein the structural backbone passes through at least one rotor systemand connects the rotor inflow side to the rotor outflow side of therotor system.
 43. The rotary wing aircraft of claim 42, furthercomprising a central buss, and wherein the central buss is configured toconduct power and control signals from an inflow side to an outflow sideof the rotor through the structural backbone.
 44. The rotary wingaircraft of claim 43, further comprising electrical system componentsappended to the structural backbone, and wherein power and controlsignals to and between the electrical system components are conducted onthe central buss through the structural backbone.
 45. The rotary wingaircraft of claim 44, further comprising a modular aircraft componenthaving an aperture receptive to the structural backbone, and whereinelectrical connectors for connecting electrical components to thecentral buss are recessed within the structural backbone so the modularaircraft component is slideable along the length of the backbone withoutdamaging an electrical connector during one of assembly and disassemblyof the aircraft.
 46. The rotary wing aircraft of claim 42, wherein thestructural backbone is generally circular in cross section.
 47. Therotary wing aircraft of claim 42, wherein the structural backbone isgenerally cruciform in cross section.
 48. The rotary wing aircraft ofclaims 42, wherein the structural backbone is made of a fiber reinforcedplastic material such as epoxy bonded carbon fiber.
 49. On a UAV havingan airframe, a propulsion system, a control system and a payload, meansfor withstanding the launch loads produced when deploying the UAV from alaunch tube with a charge actuated device (CAD) comprising a structuralbackbone extending along a longitudinal axis of the UAV and supportingthe airframe, propulsion system, control system, and payload.
 50. Amethod of munition damage assessment, the method comprising the stepsof: equipping a munition with an electrically powered rotary wing UAVhaving a sensor such as a video camera and a telemetry system,delivering the UAV to the vicinity of a target site concurrently withthe munition, commanding the UAV to hover in the vicinity of the targetsite to observe damage caused by the munition, and transmittinginformation from the UAV to a remote location through a telemetry link.51. The method of claim 50, the method further comprising the steps ofattaching the UAV to the munition and delivering the UAV to the vicinityof the target site by the munition and releasing the UAV from themunition before it reaches the target site allowing the UAV to orbit inthe vicinity of the target site and observe damage caused by themunition.